1. Field of the Invention
The present invention relates to systems and methods for spacecraft navigation and control, and in particular to a system and method for determining spacecraft momentum and for dumping that momentum when necessary.
2. Description of the Related Art
Almost all Earth orbiting spacecraft utilize a set of reaction wheels to control attitude. A reaction wheel is a continuously spinning flywheel, mounted on a bearing. Most spacecraft use three reaction wheels to control attitude, with a fourth redundant wheel available in case of failure. As is the case with any spinning object, a force (or torque) can be created by altering the rotation rate of a reaction wheel. By carefully controlling the rotational speeds of the individual reaction wheels, the resulting torques can be combined to rotate the spacecraft to the desired orientation in space.
If a reaction wheel were to ever stop spinning, it would create an abrupt change in torque that would cause the spacecraft to jitter. This sudden jitter results in undesirable attitude errors. Therefore, reaction wheels are kept continuously spinning. The energy required to speed up and slow down the spinning reaction wheels can be collected from the Sun via the spacecraft's solar panels, or obtained from other sources.
Due to physical or material constraints, continuous-spin reaction wheel mechanisms are built to handle only up to a particular spin speed. After that speed is achieved, they can be accelerated no more. It is then necessary to “dump” the momentum built up in the reaction wheels. Typically, the spacecraft's momentum management software continuously makes adjustments to control momentum buildup.
Momentum can be dumped using thrusters, torquer bars, or similar devices. These devices create a torque external to the spacecraft opposing the wheel momentum direction to dump the wheel momentum. Momentum management software employs these devices to compensate for secular (non-periodic) momentum increases, allowing it to maintain reaction wheel spin rate within the designed operational range.
To determine when to dump momentum from the spacecraft and how much to dump, an estimate of the spacecraft's momentum must be made. Current estimators are torque-based in that they operate by estimating torque, then convert the estimated torque to equivalent momentum accumulation. This is accomplished by accumulating momentum changes (including that of the central body, reaction wheels, solar wings, and other appendages) over a period of about a minute. A backwards difference of momentum is taken over time as a measured torque, and used to update torque estimates. Torque is typically modeled as a Fourier series including a constant and two sinusoidal terms.
The torque estimate is separated into secular and cyclic estimates, typically in an earth centered inertial (ECI) reference frame. Secular momentum, which grows over time, is estimated and dumped on a periodic basis to avoid saturating the momentum wheels. The dumped momentum is a combination of the secular momentum estimate and the product of the secular torque estimate and one sidereal day. Typically, the momentum estimator is suspended when there is a thruster or engine firing, so that such control torques will not be mistaken as environmental torques.
Unfortunately, since environmental torques are not sinusoidal when transitioning to the eclipse (when the satellite enters and leaves the Earth's shadow), the use of a second order Fourier series does not provide sufficient accuracy to approximate the environmental torque, and inaccurate torque estimates provide inaccurate momentum dumping.
Also, the process of backwards differencing momentum over time to arrive at a measured torque is inherently noisy, and can result in substantial noise amplification. This noise amplification makes it difficult for a torque estimate to converge. Further, the foregoing technique is characterized by a long time constant and large phase large that results in a low stability margin in the momentum dumping control loop.
The foregoing techniques fail to use a momentum repeatable period of a solar day, as the torque estimator is designed to be repeatable in a sidereal day to accommodate satellite stationkeeping. This results in a momentum walking of four minutes per day, and further degrades performance. The foregoing estimation technique also dumps small quantities of momentum when it might be tolerable to save such small quantities for the next scheduled momentum dump. Further, the estimator sampling/updating period cannot be adjusted on orbit.
What is needed is a method and apparatus for estimating spacecraft momentum and for dumping excessive momentum without the foregoing disadvantages. The present invention satisfies this need.